Thermal management of a gas turbine engine shaft

ABSTRACT

A gas turbine engine includes a fan rotor, a compressor section, a combustor section, and a turbine section. The turbine section is positioned downstream of the combustor section. A fan drive turbine in the turbine section, and a shaft connects the fan drive turbine to the fan rotor. An inlet duct is connected to a cooling air source and connected to a cooling compressor downstream of the fan drive turbine. The cooling compressor is connected to an air source, and connected to a turning duct for passing compressed air in an upstream direction through the shaft. A method is also disclosed.

BACKGROUND

This application relates to shaft cooling for a turbine driven shaft ina gas turbine engine.

Gas turbine engines are known and typically include a fan deliveringbypass air into a duct as propulsion air. Air is also delivered into acore engine where it passes into a compressor. The air is compressed anddelivered into a combustion section where it is mixed with fuel andignited. Products of this combustion pass downstream over turbinerotors, driving them to rotate. The turbine rotors drive the fan andcompressor rotors.

Historically, a turbine rotor has driven a fan rotor through a directconnection and at a common speed. However, more recently, gas turbineengines have been developed wherein a gear reduction is placed betweenthe two such that the fan rotor rotates at a slower speed than the fandrive turbine. With this change, the diameter of the fan can increase,and a bypass ratio or amount of air delivered into the bypass ductcompared to the air delivered into the core engine, can increase.

With such changes, the core engine and the turbine sections becomesmaller. In addition, the torque and heat load applied to shaftsconnecting the turbine rotors and the compressors and fan rotor alsoincrease.

SUMMARY

In a featured embodiment, a gas turbine engine includes a fan rotor, acompressor section, a combustor section, and a turbine section. Theturbine section is positioned downstream of the combustor section. A fandrive turbine in the turbine section, and a shaft connects the fan driveturbine to the fan rotor. An inlet duct is connected to a cooling airsource and connected to a cooling compressor downstream of the fan driveturbine. The cooling compressor is connected to an air source, andconnected to a turning duct for passing compressed air in an upstreamdirection through the shaft.

In another embodiment according to the previous embodiment, connectionpassages in the shaft communicates inner and outer passages in the shaftsuch that the air, having passed in the upstream location in one of theinner and outer passages is returned in the other of the inner and outerpassages in a downstream direction to provide cooling to the shaft.

In another embodiment according to any of the previous embodiments, theturning duct extends to initially pass air downstream of the coolingcompressor in a downstream direction, and then turn the air into theupstream direction.

In another embodiment according to any of the previous embodiments, atube is positioned within the shaft to define the inner and outerpassages, and the connection passages extend through the tube tocommunicate between the inner and outer passages.

In another embodiment according to any of the previous embodiments, aplurality of bearings support the shaft.

In another embodiment according to any of the previous embodiments,bearing outlet passages extend through the shaft. The bearing outletpassages are positioned to provide cooling air to the bearings.

In another embodiment according to any of the previous embodiments, theinlet duct communicates with a bypass duct that is positioned outwardlyof a core engine including the compressor section, the combustorsection, and the turbine section, and the bypass duct providing thesource of air.

In another embodiment according to any of the previous embodiments, thecooling compressor is connected to the shaft to a flexible connection.

In another embodiment according to any of the previous embodiments, theturning duct is positioned relative to a flange at an end of the shaftwith an intermediate seal to provide a sealed connection between theduct and the shaft.

In another embodiment according to any of the previous embodiments, thecooling compressor is supported by a bearing which also supports theshaft.

In another embodiment according to any of the previous embodiments, agear reduction is positioned between the fan drive turbine and the fanrotor.

In another embodiment according to any of the previous embodiments, aplurality of bearings support the shaft.

In another embodiment according to any of the previous embodiments,bearing outlet passages extend through the shaft. The bearing outletpassages are positioned to provide cooling air to the bearings.

In another embodiment according to any of the previous embodiments, aplurality of bearings support the shaft.

In another embodiment according to any of the previous embodiments,bearing outlet passages extend through the shaft. The bearing outletpassages are positioned to provide cooling air to the bearings.

In a featured embodiment, a gas turbine engine includes a fan rotor, acompressor section, a combustor section, and a turbine section. Theturbine section is positioned downstream of the combustor section. A fandrive turbine in the turbine section, and a shaft connects the fan driveturbine to the fan rotor, an outer passage within the shaft, and aninner passage within the shaft. There is a means for supplying coolingair to a cooling compressor downstream of the fan drive turbine. Thereis a means for passing compressed air in an upstream direction throughone of the inner and outer passages in the shaft, and providingcommunication between the inner and outer passages such that the air,having passed in the upstream location may be returned in the other ofthe inner and outer passages in a downstream direction to providecooling to the shaft.

In another embodiment according to the previous embodiment, the meansfor passing compressed air includes a tube positioned within the shaftto define the inner and outer passages, and connection passagesextending through the tube to communicate between the inner and outerpassages.

In another embodiment according to any of the previous embodiments, themeans for supplying cooling air includes an inlet duct that communicateswith a bypass duct that is positioned outwardly of a core engineincluding the compressor section, the combustor section, and the turbinesection, and the bypass duct providing the cooling air.

In another featured embodiment, a method of operating a gas turbineengine includes the steps of driving a fan drive turbine in a turbinesection, and driving a shaft connecting the fan drive turbine to drive afan rotor, an outer passage within the shaft, and an inner passagewithin the shaft. Cooling air is supplied to a cooling compressordownstream of the fan drive turbine, and the cooling compressor isconnected to an air source. Compressed air is passed in an upstreamdirection through one of the inner and outer passages in the shaft, andcommunicating the inner and outer passages such that the air, havingpassed in the upstream location is returned in the other of the innerand outer passages in a downstream direction to provide cooling to theshaft.

In another embodiment according to the previous embodiment, the step ofinitially passing air downstream of the cooling compressor in adownstream direction, and then turning the air into the upstreamdirection.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows a cooling scheme.

FIG. 3 shows cooling airflow paths through the FIG. 2 engine.

FIG. 4 shows further details of area 4 from FIG. 3.

FIG. 5 shows another potential feature.

FIG. 6 shows an alternative embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle15, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 shows a cooling scheme for engine 100. A combustor 102 is shownalong with a high pressure turbine 104 having a shaft 105 rotatingoutwardly of a low pressure spool and shaft 138.

A fan drive or low pressure turbine 106 rotates with the shaft 138.Exhaust airflow 108 is shown downstream of the fan drive turbine 106. Aworker of ordinary skill in this art would recognize that the term“downstream” would mean passing from the compressor (see FIG. 1) throughthe combustor, across the turbine section, and to the exhaust such asshown at core airflow 108. Upstream is in the opposed direction.

A duct inlet 112 is shown extending into a bypass duct 110. The inlet112 communicates with the duct 114 passing through a turbine exhaustcase 115.

This air passes across a vane 116 and to a cooling compressor 118. Thecooling compressor delivers the airflow across a vane 122 and into adischarge duct 120 that takes the airflow initially in a downstreamdirection, but then turns the cooling airflow at elbow 121 back in anupstream direction. As shown, the duct 120 has end flanges 126 and 128capturing an end 134 of a downstream shaft portion 132. A seal 130 sitsbetween the portion 134 and the flange 126.

The downstream shaft portion 132 rotates with the shaft 138. Downstreamportion 132 may be a flexible input coupling as shown with theundulations at 133. The shaft portion 132 drives the cooling compressor118. Thus, the air passing into the duct 120 and ultimately into apassage 136, in the shaft portion 132 and the shaft 138, is at a higherpressure than that tapped from the bypass duct 110. As shown, a returntube 140 is positioned within the passage 136. As will be explained, airpasses from the elbow 121, upstream through the passage 136, and thenradially inwardly into the tube 140 at an upstream location, and thenpasses back downstream through a tube portion 142. Tube portion 142 isshown venting at 143 at a location associated with the tail cone 144.

With this arrangement, the shaft portion 132 and the shaft 138 areprovided with high pressure cooling air that is able to pass along thetorturous flow path, as described, and provide adequate cooling.

In addition, bearings 150 and 152 are illustrated along with bearings154 and 156. Bearings 150 and 152 support shaft 138. The bearings 154and 156 support the shaft portion 132.

FIG. 3 shows the same basic structure, however, now including arrowswhich show the cooling air path. Heat passes into the shaft at locationsassociated with the bearings and the turbine rotors. As can beappreciated, the cooling scheme provided in FIGS. 2 and 3 will provideadequate cooling to those areas.

FIG. 4 shows a further detail of area 4 from FIG. 3. As shown, the tube140 has passages 151 allowing the cooling air passing upstream to thepassage 136 to turn radially inwardly into the tube 140 through openings151 into the interior 153 of the tube and pass to downstream locations142.

FIG. 5 shows an alternative feature wherein the shafts 132/138 outwardlyof the passage 136 may be provided with a plurality of cooling holes 170to provide cooling to the location of the bearings 150, 152, 154, 156.By providing bearing cooling airflow at the high pressures achieved withthe cooling compressor 118, buffering and thermal conditioning of thebearings will be provided.

FIG. 6 shows an alternative embodiment 200 wherein the duct 214 suppliesair across a vane 254, a cooling compressor 252, a vane 256, and intothe duct 224. Air in the duct 224 will communicate with a passage 236and will provide cooling to the shaft as in the above embodiment. An end250 of the rotor is provided adjacent a seal 251 and a seal 261 sealingon a downstream end 260 of the shaft. Seal 261 may be a brush seal,although other type seals are contemplated. By locating the compressorrotor radially inwardly, the rear bearings 256 are associated not onlywith the rear shaft 232, but also with the compressor rotor 252.

One could define a gas turbine engine coming within this disclosure ashaving a fan rotor, a compressor section, a combustor section, and aturbine section. The turbine section is positioned downstream of thecombustor section. There is a fan drive turbine in the turbine section,and a shaft connects the fan drive turbine to the fan rotor. There isouter passage within the shaft, and an inner passage within the shaft.There is a means for supplying cooling air to a cooling compressordownstream of the fan drive turbine. There is also a means for passingcompressed air in an upstream direction through one of the inner andouter passages in the shaft, and communicate the inner and outerpassages such that the air, having passed in the upstream location maybe returned in the other of the inner and outer passages in a downstreamdirection to provide cooling to the shaft.

A method of operating a gas turbine engine includes the steps of drivinga fan drive turbine in a turbine section. Also, driving a shaftconnecting the fan drive turbine to drive a fan rotor, an outer passagewithin the shaft, and an inner passage within the shaft. Supplyingcooling air to a cooling compressor downstream of the fan drive turbine.Suppling compressed air in an upstream direction through one of theinner and outer passages in the shaft. The inner and outer passagescommunicate such that the air, having passed in the upstream location isreturned in the other of the inner and outer passages in a downstreamdirection to provide cooling to the shaft.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this disclosure. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this disclosure.

The invention claimed is:
 1. A gas turbine engine comprising: a fanrotor, a compressor section, a combustor section, and a turbine section,said turbine section being positioned downstream of said combustorsection; a fan drive turbine in said turbine section, and a shaftconnecting said fan drive turbine to said fan rotor, said shaft alsodriving a cooling compressor; an inlet duct connected to a cooling airsource and connected to provide air to said cooling compressordownstream of said fan drive turbine, and said cooling compressor isconnected to a turning duct for passing compressed air from said coolingcompressor in an upstream direction through said shaft; and whereinthere being connection passages in said shaft to provide communicationbetween inner and outer passages in said shaft such that the compressedair, having passed in said upstream direction through said shaft in oneof said inner and outer passages is returned in the other of said innerand outer passages in a downstream direction to provide cooling to saidshaft.
 2. The gas turbine engine as set forth in claim 1, wherein saidturning duct extending to initially pass the compressed air downstreamof said cooling compressor in said downstream direction, and then turnthe compressed air into the upstream direction.
 3. The gas turbineengine as set forth in claim 2, wherein a tube is positioned within saidshaft to define said inner and outer passages, and said connectionpassages extend through said tube to communicate between said inner andouter passages.
 4. The gas turbine engine as set forth in claim 3,wherein a plurality of bearings support said shaft.
 5. The gas turbineengine as set forth in claim 4, wherein bearing outlet passages extendthrough said shaft, said bearing outlet passages positioned to providecooling air to said bearings.
 6. The gas turbine engine as set forth inclaim 1, wherein said inlet duct communicates with a bypass duct that ispositioned outwardly of a core engine including said compressor section,said combustor section, and said turbine section, and said bypass ductproviding said cooling air source.
 7. The gas turbine engine as setforth in claim 1, wherein said cooling compressor is connected to saidshaft by a flexible shaft portion having undulations.
 8. The gas turbineengine as set forth in claim 1, wherein a gear reduction is positionedbetween said fan drive turbine and said fan rotor.
 9. The gas turbineengine as set forth in claim 8, wherein a plurality of bearings supportsaid shaft.
 10. The gas turbine engine as set forth in claim 9, whereinbearing outlet passages extend through said shaft, said bearing outletpassages positioned to provide cooling air to said bearings.
 11. The gasturbine engine as set forth in claim 1, wherein a plurality of bearingssupport said shaft.
 12. The gas turbine engine as set forth in claim 11,wherein bearing outlet passages extend through said shaft, said bearingoutlet passages positioned to provide cooling air to said bearings. 13.A method of operating a gas turbine engine comprising the steps of:driving a fan drive turbine in a turbine section, and driving a shaftconnecting said fan drive turbine to drive a fan rotor, an outer passagewithin said shaft, and an inner passage within said shaft; supplyingcooling air into an inlet duct and then to a cooling compressor drivenby said shaft downstream of said fan drive turbine, and passingcompressed air from the cooling compressor in an upstream directionthrough one of said inner and outer passages in said shaft, andproviding communication between said inner and outer passages such thatthe compressed air, having passed in said upstream direction is returnedin the other of said inner and outer passages in a downstream directionto provide cooling to said shaft; and said one of said inner and outerpassages being said outer passage, with the other of said inner andouter passages being said inner passage.
 14. The method as set forth inclaim 13, including the step of initially passing the compressed airdownstream of said cooling compressor in a downstream direction, andthen turning the compressed air into the upstream direction.
 15. A gasturbine engine comprising: a fan rotor, a compressor section, acombustor section, and a turbine section, said turbine section beingpositioned downstream of said combustor section; a fan drive turbine insaid turbine section, and a shaft connecting said fan drive turbine tosaid fan rotor, said shaft also driving a cooling compressor; an inletduct connected to a cooling air source and connected to provide air tosaid cooling compressor downstream of said fan drive turbine, and saidcooling compressor is connected to a turning duct for passing compressedair from said cooling compressor in an upstream direction through saidshaft; connection passages in said shaft to provide communicationbetween inner and outer passages in said shaft such that the compressedair, having passed in said upstream direction through said shaft in oneof said inner and outer passages is returned in the other of said innerand outer passages in a downstream direction to provide cooling to saidshaft; said turning duct extending to initially pass the compressed airdownstream of said cooling compressor in said downstream direction, andthen turn the compressed air into the upstream direction; and wherein atube is positioned within said shaft to define said inner and outerpassages, and said connection passages extend through said tube tocommunicate between said inner and outer passages, with compressed airdownstream of said turning duct intially passing through said outerpassage in said upstream direction, and then turning into said innerpassage to return in said downstream direction.
 16. The gas turbineengine as set forth in claim 15, said turning duct has a downstream endwith an end flange capturing an end of said shaft, with a seal sittingbetween said end flange and said shaft.
 17. The gas turbine engine asset forth in claim 1, wherein said turning duct has a downstream endwith an end flange capturing an end of said shaft, with a seal sittingbetween said end flange and said shaft.